First equipment heating device

ABSTRACT

An aircraft provided with at least two pieces of aeronautic equipment, a first piece of equipment ( 25 ) including a part intended to be arranged at a skin ( 27 ) of the aircraft and elements for heating the part, characterized in that the heating elements include a thermodynamic loop including a closed circuit in which a heat transfer fluid circulates, the closed circuit including an evaporator ( 14 ) associated with functional elements ( 80 ) of the second piece of equipment ( 81 ) of the aircraft forming a heat source giving off heat during their operation and a zone in which a condensation of the heat transfer fluid can occur in the appendage to heat it, and in that outside the evaporator ( 14 ), the circuit in which the fluid circulates is formed by a tubular channel with an empty section.

The invention relates to an aircraft, and more particularly to theheating of a first piece of aeronautic equipment intended to be arrangedat the skin of the aircraft.

To perform its mission, an aircraft comprises several pieces ofequipment comprising parts flush with or appendages protruding from theskin of the aircraft.

These appendages or these flush parts for example belong to probes inparticular making it possible to measure different aerodynamicparameters of the airflow surrounding the aircraft, in particular thetotal pressure, static pressure, temperature, or incidence of theairflow near the skin of the aircraft.

The total pressure, combined with the static pressure, makes it possibleto determine the local speed of the airflow near the probe.

Other probes for example make it possible to measure the local incidenceof an airflow.

The incidence probes may comprise moving appendages intended to beoriented in the axis of the airflow surrounding the probe.

The orientation of the probe makes it possible to determine theincidence of the airflow.

Other incidence probes may be equipped with stationary appendagesequipped with several pressure taps.

The pressure difference measured between these pressure taps makes itpossible to determine the incidence of the airflow surrounding theprobe.

Other pieces of equipment such as cameras also must be installed flushor protruding relative to the skin of the aircraft, for example on pods.

During flight at high altitudes, the aircraft may encounter freezingconditions.

More specifically, ice may form on the skin and appendages of theaircraft. The appearance of ice is particularly problematic for theaerodynamic probes, the profiles of which may be modified by ice and thepressure tap orifices of which may be obstructed.

The measuring instruments mounted on pods may also be disrupted by theappearance of ice.

One solution making it possible to avoid ice formation consists ofheating the appendages.

Currently, in most cases, heating is done using electrical resistancesembedded in the appendages.

The heating is done by Joule effect. For example, to heat a totalpressure probe, it is necessary to dissipate several hundred watts.

More specifically, this type of probe is formed by a mast bearing a tubeclosed at one of its ends and called Pitot tube.

The heating of the probe is done using a heating resistance made in theform of a heating wire wound around the body of the probe, i.e., both inthe mast and the Pitot tube.

To produce the heating wire, an electrical conductor is commonly usedincluding an alloy of iron and nickel coated with an insulating materialsuch as alumina or manganese. The insulator itself is coated with anickel or Inconel sheath allowing brazing of the wire on the body of theprobe.

A method for producing such a probe is for example described in patentapplication FR 2,833,347 filed in the applicant's name.

The production of the heating wire and its assembly in the probe requirea series of complex and costly operations.

Another embodiment to heat a Pitot tube probe had been considered inU.S. Pat. No. 4,275,603.

This document describes the use of a heat pipe contributing heat energyaround the tube. The return of the heat transfer fluid to liquid stateis ensured in a porous material.

This allows the probe to be arranged in any possible orientation on theskin of the aircraft.

In practice, this solution has no industrial advantage due to thedifficulty of inserting a porous material in a probe.

The method for producing such a probe is at least as complex as thatimplementing a heating wire.

The invention seeks to propose a new heated probe, and more generally apiece of aeronautic equipment that is flush or that has a heated outerappendage, the production of which is much simpler than that describedin the prior art.

To that end, the invention relates to an aircraft provided with at leasttwo pieces of aeronautic equipment, a first piece of equipmentcomprising at least one part intended to be arranged at a skin of theaircraft and means for heating the part, characterized in that theheating means comprise a thermodynamic loop comprising a closed circuitin which a heat transfer fluid circulates, the closed circuit comprisingan evaporator associated with functional means of the second piece ofequipment of the aircraft forming a heat source giving off heat duringtheir operation and a zone in which a condensation of the heat transferfluid can occur in the appendage to heat it, and in that outside theevaporator, the circuit in which the fluid circulates is formed by atubular channel with an empty section.

According to other features of the aircraft according to the invention,considered alone or in combination:

-   -   the functional means of the second piece of equipment of the        aircraft forming a heat source are formed by a part or an        electronic board thereof;    -   the functional means of the second piece of equipment of the        aircraft forming a heat source are formed by an actuator        thereof;    -   the functional means of the second piece of equipment of the        aircraft forming a heat source are formed by a part or a piece        of electrical equipment thereof;    -   the channel is configured for the fluid to circulate therein by        capillarity;    -   it comprises a circulation pump for the heat transfer fluid;    -   the tubular channel forms a single thermodynamic loop outside        the evaporator;    -   the tubular channel forms several thermodynamic loops in which        the heat transfer fluid circulates in parallel outside the        evaporator;    -   the part is configured to be flush with the skin of the        aircraft;    -   the part is an appendage configured to be arranged protruding        relative to the skin of the aircraft;    -   the first piece of equipment comprises a base intended to fasten        the piece of equipment on the skin of the aircraft, the        appendage is arranged on a first side of the base and the        evaporator is arranged on a second side of the base, opposite        the first side;    -   the first piece of equipment comprises an aerodynamic measuring        probe.

According to another aspect, the invention also relates to a method forproducing a first piece of aeronautic equipment of an aircraft aspreviously described, the equipment comprising a body in which thetubular channel with an empty section is produced, the method beingcharacterized in that the body is produced using an additivemanufacturing method.

Lastly, the invention also relates to a data file stored on storagemeans and able to be loaded in the memory of a processing unitassociated with an additive manufacturing machine able to manufacture anobject by superimposing layers of material, characterized in that itcomprises data for three-dimensional depiction of the first piece ofequipment for an aircraft as previously described, so as to allow, whenit is loaded into memory of, and processed by, said processing unit, themanufacture of said piece of equipment by said additive manufacturingmachine.

The invention will be better understood, and other advantages thereofwill appear, upon reading the detailed description of one embodimentgiven as an example, this description being illustrated by the attacheddrawing, in which:

FIG. 1a diagrammatically shows a thermodynamic loop able to heat a pieceof aeronautic equipment;

FIG. 1b diagrammatically shows several thermodynamic loops able to heata piece of aeronautic equipment;

FIG. 2 shows an aerodynamic probe intended to measure the total pressureand equipping an aircraft;

FIGS. 3a and 3b show a mast and a Pitot tube forming outer parts of theprobe of FIG. 1;

FIG. 4 shows an exploded view of different component parts of the probe;

FIGS. 5a and 5b show an aerodynamic probe intended to measure the staticpressure and equipping an aircraft;

FIG. 6 shows an example embodiment of an aerodynamic probe included inthe makeup of an aircraft according to the invention.

For clarity reasons, the same elements bear the same references in thedifferent figures.

FIG. 1a diagrammatically shows a thermodynamic loop 11 in which a heattransfer fluid circulates in a closed circuit.

In this loop, the fluid may assume two phases: liquid 12 and vapor 13.

The latent transformation heat between these two phases is used totransport heat energy between an evaporator 14 and a condenser 15.

This type of thermodynamic loop is widely used to cool electroniccomponents dissipating heat during their operation.

In general, a heat contribution, diagrammed by arrows 16, at theevaporator 14, is transported by the fluid in vapor phase 13 toward thecondenser 15, where the energy contribution is returned to thesurrounding environment.

This return is diagrammed by arrows 17.

The closed circuit also comprises a reservoir 18 containing heattransfer fluid in liquid state. The reservoir 18 is arranged near theevaporator 14. The reservoir 18 supplies the loop 11 via the evaporator14.

Thus, once a sufficient energy contribution is captured by theevaporator 14, the fluid in liquid state contained in the evaporatorvaporizes. The overpressure due to the evaporation pushes the fluid invapor state 13 toward the condenser 15, where the fluid regains itsliquid state to return toward the evaporator 14.

In the present application, the thermodynamic loop 11 is used to heatpart of a piece of onboard aeronautic equipment.

Onboard an aircraft, many pieces of equipment have protruding appendagesrelative to the skin of the aircraft or flush parts.

These pieces of equipment may be aerodynamic probes, antennas, sensors,etc.

These appendages or these flush parts require heating to allow them tooperate. This heating is particularly important for the aerodynamicprobes, which have orifices used as pressure taps.

The heating makes it possible to avoid the formation of ice, which couldobstruct these orifices.

The fire probes, which have a vane intended to be oriented in the bed ofthe airflow surrounding the probe, are also sensitive to the ice thatmay form on the vane and alter its shape, thus causing an incorrectmeasurement, or even blocking of the vane.

FIG. 1b diagrammatically shows two thermodynamic loops 11 a and 11 b inwhich the heat transfer fluid circulates in parallel outside anevaporator 14 shared by the different loops.

These different loops 11 a and 11 b more specifically make it possibleto heat different zones, forming condensers 15 a and 15 b, of anappendage or part of a piece of aeronautic equipment.

The invention may of course be implemented for more than twothermodynamic loops.

FIG. 2 shows an aeronautic probe 25 making it possible to measure thetotal pressure of an airflow surrounding the skin 27 of an aircraft.

The probe 25 is intended to be fixed crossing through an opening 26formed in the skin 27 of the aircraft.

In FIG. 1, the skin 27, at its opening 26, is shown at a distance fromthe probe 25.

The probe 25 comprises a Pitot tube 30 and a mast 31 supporting thePitot tube 30.

The Pitot tube 30 and the mast 31 are outside the skin 27.

The probe 25 also comprises a part inside the skin 27 including apneumatic connector 32 allowing the pneumatic connection of the Pitottube 30 to a pressure sensor situated inside the fuselage of theaircraft.

The probe 25 is positioned on the skin 27 of the aircraft such that thePitot tube 30 is oriented substantially along a longitudinal axis of theaircraft, outside the boundary layer, so that the direction of the flow,embodied by an arrow 33, is substantially across from an inlet orifice34 situated at first end 35 of the Pitot tube 30.

A second end 36 of the Pitot tube 30, opposite the end 35, is closed soas to create a stop point in the air taken from the flow and penetratingthe tube 30 through its orifice 34.

At the end 36 of the tube, a pneumatic channel, not shown in FIG. 1,opens in the tube 30 to form a pressure tap therein at which one seeksto measure the air pressure.

The pneumatic channel is for example connected to a pressure sensor oranother pressure measuring device, for example a flowmeter.

The pressure sensor allows an effective measurement of the air pressureprevailing inside the tube 30 at its obstructed end 36.

The pressure sensor can belong to the probe 25 or be offset. In thiscase, the pressure sensor is connected to the probe 25 using a hose andthe pneumatic connector 32.

At the end 36, the tube 30 includes one or several drain holes, notshown, allowing the discharge of the water penetrating the inside of thetube 30.

Aside from the bleed hole(s), which have a small section relative tothat of the tube 30, the tube 30 is closed at its end 36.

The pressure measured at this end therefore represents the totalpressure Pt of the flow of air.

The mast 31 bears the Pitot tube 30 at its second end 36.

The Pitot tube 30 has a substantially cylindrical shape and the mast 31has an elongated shape. The mast 31 is for example in the shape of awing, the concave and convex sides of which may be symmetrical.

The probe 25 may comprise other pressure taps, for example pressure tapsarranged on the mast 31 or around the tube 30 on its cylindrical partand making it possible to define the local incidence of the flowrelative to the probe 25 or measuring the static pressure of the flow.

The probe 25 comprises fastening means intended to fasten the probe 25to the skin 27 of the aircraft.

These means for example comprise a base 38 formed by a shoulder intendedto come into contact with the skin 27.

Screws arranged around the opening 26 immobilize the base 38 relative tothe skin 27.

In the illustrated example, the Pitot tube 30 is stationary relative tothe skin 27 of the aircraft.

It is of course possible to mount the Pitot tube 30 on a moving mast,for example a vane that may be oriented in the axis of the flow, as forexample described in the patent published under no. FR 2,665,539 andfiled on Aug. 3, 1990.

The base 38 then comprises a pivot link allowing the rotation of themast 31 relative to the skin 27 around an axis perpendicular to the skin27.

Thus, when the local incidence of the flow, near the probe 25, evolves,the orientation of the Pitot tube 30 follows this incidence so as alwaysto face the flow.

The total pressure measurement Pt is thereby improved during localincidence variations of the flow along the skin 27 of the aircraft.

The evaporator 14 and the reservoir 18 are arranged inside the fuselageof the aircraft on one side of the base 38.

The condenser 15 is formed by a channel arranged in the mast 31 and inthe Pitot tube 30.

Heating means make it possible to contribute heat energy to theevaporator 14.

These means for example comprise a heating electrical resistance 40arranged around the evaporator 14.

Any other means making it possible to contribute heat to the evaporatormay also be implemented in the context of the invention, for example thepassage of a hot air flow along the outer walls of the evaporator 14.

Of course, other means may be considered, as will be described in moredetail below.

It is possible to place a temperature sensor in the appendage, making itpossible to measure its temperature to enslave the heating means.

Alternatively, a temperature measurement of the fluid in the evaporator14 provides an image of the temperature of the appendage.

Using a thermodynamic loop to heat the probe 25, and more generally anaeronautic appendage, has the advantage of facilitating the regulationof the temperature of the appendage by controlling the heating meansdelocalized inside the skin of the aircraft near the appendage.

The fluids generally used as heat transfer fluids in a diphasicthermodynamic loop can have high latent transformation heats, whichmakes it possible to reduce the fluid flow rate in the loop for a sameheat exchange.

The reduction in flow rate makes it possible to reduce the pressurelosses in the loop.

As an example, methanol may be used as heat transfer fluid.

In the situation described above, the fluid circulates in a tubularchannel 39 with an empty section between the evaporator 14 and thecondenser 15, in the condenser 15 itself, and between the condenser 14and the evaporator 14.

In other words, outside the evaporator 14, the circuit in which thefluid circulates is formed by the tubular channel 39 with an emptysection.

A tubular channel with an empty section refers to a channel notincluding any filling, other than the fluid, of course.

In particular, no porous material is present in the tubular channel 39.The inner walls of the tubular channel 39 are smooth to facilitate thecirculation of the fluid and limit the pressure losses.

FIGS. 3a and 3b show an example arrangement of the channel 39 equippingthe outer parts of the probe 25 and in which the heat transfer fluidcirculates that makes it possible to heat these outer parts.

The mast 31 and the tube 30 both comprise an enclosure, 41 for the mast,and 42 for the tube 30.

The pneumatic channel used for the pressure measurement circulates inthe enclosure 41. The channel 39 is made in the respective enclosures.

In the channel 39, the fluid that circulates therein is able to condenseto heat the corresponding enclosure or part of that enclosure as needed.

More specifically, another advantage related to the production of thetubular channel 39 with an empty section is the auto-adaptation capacityof the heat exchanges at the probe.

Indeed, the exchange coefficient between the fluid and the wall,condensation coefficient, is related to the temperature gradientsbetween the fluid and the wall.

The heat exchanges are greater in the coldest zones of the probe 25.These coldest zones correspond to the zones of the enclosures where theouter cooling is greatest.

This makes it possible to obtain better homogeneity of the probe interms of temperature.

FIG. 3a shows the mast 31 and the pitot tube 30 in profile. One examplepath of the channel 39 in the corresponding enclosures can be seen inthis figure.

FIG. 3b shows the mast 31 in sectional view in a plane parallel to theskin 27 near the opening 26.

Along its path, the channel 39 can be broken down into three parts 39 a,39 b and 39 c, following one another.

After it leaves the evaporator 14, the fluid circulates in the part 39 amade in the enclosure 41. The part 39 a can wind in the enclosure 41between the leading edge 31 a and the trailing edge 31 b of the mast 31.

Next, the part 39 b of the channel 39 winds in the enclosure 42. Thepath of the part 39 b is for example helical around the inner cavity ofthe Pitot tube 30 at the bottom of which the total pressure is measured.

The channel 39 continues its path in the part 39 c while againcirculating in the enclosure 41 of the mast 31.

As for the part 39 a, the part 39 c can wind in the enclosure 41 betweenthe leading edge 31 a and the trailing edge 31 b of the mast 31.

The definition of the path of the channel 39 is done based on the zonesof the probe that should preferably be heated.

In the illustrated example, the channel 39 winds in the appendage whileforming a single loop outside the evaporator 14.

It is also possible to produce several thermodynamic loops in theappendage, in which loops the heat transfer fluid circulates in paralleloutside the evaporator 14, as shown diagrammatically in FIG. 1 b.

The auto-adaptation of the heat exchange to the actual temperature ofthe outer walls of the probe 25 allows a more tolerant definition of thepath than for a probe heated directly by an electrical resistance.

The section of the channel may vary along its entire path in the mast 31and in the Pitot tube 30.

The circulation of the fluid in the channel 39 may be ensured using acirculation pump 45 arranged upstream from the evaporator 14. Thecirculation pump 45 is advantageously arranged inside the skin 27 of theaircraft.

Alternatively, it is possible to do away with this circulation pump 45by configuring the section of the different parts 39 a to 39 c of thechannel 39 so that the fluid circulates in its liquid phase bycapillarity.

Such a circulation mode requires relatively small sections.

In order to retain a sufficient overall flow rate, the channel 39 maycomprise zones placed in parallel.

It is advantageous to produce the probe 25, and more generally any pieceof aeronautic equipment implementing the invention, by carrying out anadditive manufacturing method to manufacture the mechanical part(s) inwhich the channel 39 travels.

This method is also known as 3D printing.

At this time, it is known to produce metal parts using this method. Itis for example possible to use titanium-based alloys, aluminum-basedalloys, or more generally, stainless steel alloys with a base of steel,nickel and/or chromium.

FIG. 4 shows an exploded view of several mechanical parts which, whenassembled, form the probe 25.

A body 47 forms the base 38 and the enclosures 41 and 42. The channel 39can be made directly in the body 47 by additive manufacturing.

The body 47 can remain open at its trailing edge, for example toarrange, in the body, the pneumatic channels making it possible tomeasure the total pressure.

Alternatively, these channels may also be made using the additivemanufacturing method.

The trailing edge 31 a of the mast 31 and the end 36 of the Pitot tubecan be closed using a stopper 48 that can be made using any type ofmanufacturing method.

The shapes of the stopper 48 are simpler than those of the body 47. Itis for example possible to produce the stopper 48 by molding.

The additive manufacturing can of course also be used for the stopper48.

A support 49 can complete the probe 25.

The support 49 can be used to support the pneumatic connector in a firstpart 49 a as well as the evaporator 14 in a second part 49 b.

The support is assembled to the body 47 by the base 38.

FIGS. 5a and 5b show another aerodynamic probe 60.

More specifically, the probe 60 forms a piece of aeronautic equipmentcomprising a part 61 intended to be flush with the skin 27 of theaircraft.

FIG. 5a is a view in the plane of the skin 27 near the probe 60.

FIG. 5b is a sectional view perpendicular to the plane of the skin 27.

The part 61 is for example in the shape of a disc closing off an orifice62 of the skin 27. The orifice 62 is provided to receive the part 61that is fastened by screwing on the skin 27.

The probe 60 is for example a static pressure probe having one or morepressure taps 63 formed from channels emerging substantiallyperpendicular to the skin 27.

The channel 39 circulates in the part 61. The channel winds aroundpressure taps 63 in order to heat the part 61 and prevent the pressuretaps from being closed off [by] ice.

In this embodiment, the channel 39 can also form a single loop orseveral parallel loops outside the evaporator 14.

The probe 60 also comprises a part 65 inside the skin 27. The inner part65 makes it possible to receive a pressure sensor connected to thepressure taps in order to measure the static pressure of the air flowingalong the skin 27. The inner part 65 can also accommodate the evaporator14 and the reservoir 18.

Like for the probe 25, the part 61 is advantageously made by carryingout an additive manufacturing method.

The invention also relates to a data file stored on storage means andable to be loaded in the memory of a processing unit associated with anadditive manufacturing machine able to manufacture an object bysuperimposing layers of material, which comprises three-dimensionaldepiction data of the piece of equipment as previously described, so asto allow, when it is loaded in the memory of, and processed by, saidprocessing unit, the manufacture of said piece of equipment by saidadditive manufacturing machine.

FIG. 6 shows a probe 25 fastened to the skin 27 of an aircraft.

In this example embodiment, the evaporator 14 of the heating means isassociated with functional means 80 of a second functional piece ofequipment 81 of the aircraft forming a heat source for example givingoff heat through loss of the heat during their operation.

Thus for example, the functional means of the second functional piece ofequipment 81 of the aircraft can be formed by a part of or an electronicboard thereof making it possible to contribute heat energy to theevaporator, as illustrated in this FIG. 6.

Of course, still other embodiments may be considered, and these meansmay be formed by an actuator of the second functional piece of equipment81 of the aircraft or a part of or a piece of electrical equipmentthereof.

Still other embodiments may be considered.

1. An aircraft provided with at least two pieces of aeronauticequipment, a first piece of equipment comprising a part intended to bearranged at a skin of the aircraft and means for heating the part,characterized in that the heating means comprise a thermodynamic loopcomprising a closed circuit in which a heat transfer fluid circulates,the closed circuit comprising an evaporator associated with functionalmeans of the second piece of equipment of the aircraft forming a heatsource giving off heat during their operation and a zone in which acondensation of the heat transfer fluid can occur in the appendage toheat it, and in that outside the evaporator, the circuit in which thefluid circulates is formed by a tubular channel with an empty section.2. The aircraft according to claim 1, characterized in that thefunctional means of the second piece of equipment of the aircraftforming a heat source are formed by a part or an electronic boardthereof.
 3. The aircraft according to claim 1, characterized in that thefunctional means of the second piece of equipment of the aircraftforming a heat source are formed by an actuator thereof.
 4. The aircraftaccording to claim 1, characterized in that the functional means of thesecond piece of equipment of the aircraft forming a heat source areformed by a part or a piece of electrical equipment thereof.
 5. Theaircraft according to claim 1, characterized in that the channel isconfigured for the fluid to circulate therein by capillarity.
 6. Theaircraft according to claim 5, characterized in that it comprises acirculation pump for the heat transfer fluid.
 7. The aircraft accordingto claim 1, characterized in that the tubular channel forms a singlethermodynamic loop outside the evaporator.
 8. The aircraft according toclaim 1, characterized in that the tubular channel forms severalthermodynamic loops in which the heat transfer fluid circulates inparallel outside the evaporator.
 9. The aircraft according to claim 1,characterized in that the part is configured to be flush with the skinof the aircraft.
 10. The aircraft according to claim 1, characterized inthat the part is an appendage configured to be arranged protrudingrelative to the skin of the aircraft.
 11. The aircraft according toclaim 10, characterized in that the first piece of equipment comprises abase intended to fasten the piece of equipment on the skin of theaircraft, in that the appendage is arranged on a first side of the baseand in that the evaporator is arranged on a second side of the base,opposite the first side.
 12. The aircraft according to claim 1,characterized in that the first piece of equipment comprises anaerodynamic measuring probe.
 13. A method for producing a first piece ofaeronautic equipment for an aircraft according to claim 1, the equipmentcomprising a body in which the tubular channel with an empty section isproduced, the method being characterized in that the body is producedusing an additive manufacturing method.
 14. A data file stored onstorage means and able to be loaded in the memory of a processing unitassociated with an additive manufacturing machine able to manufacture anobject by superimposing layers of material, characterized in that itcomprises data for three-dimensional depiction of the piece of equipmentfor an aircraft according to claim 1, so as to allow, when it is loadedinto memory of, and processed by, said processing unit, the manufactureof said piece of equipment by said additive manufacturing machine.